While some satellite missions, such as the current global positioning system (GPS), have significantly relaxed pointing requirements, other missions, such as future GPS (GPS III) and satellite radio, require much stricter pointing requirements. These next-generation satellite missions will use yaw-steering spacecraft that operate in highly inclined orbits and that implement a Sun Nadir Pointing (SNP) attitude profile, for which the yaw attitude varies continuously throughout the orbit by as much as 180°.
The advantage of an SNP attitude profile is that it allows good power and thermal conditions to be maintained while using a standard Geosynchronous Orbit (GEO) spacecraft configuration. One disadvantage with using SNP in a highly-inclined orbit is that, because the spacecraft experiences a large change in yaw attitude, the use of a standard attitude determination system (ADS) results in degraded pointing accuracy by an approximate factor of ten due to uncompensated gyro scale factor and misalignment uncertainties.
Typical attitude determination systems (ADS), such as those used on GEO spacecraft, do not directly consider gyro scale factor and misalignment errors. These systems often include a six-state Kalman filter, with three attitude states and with three gyro bias states that are assumed to be slowly varying constants. For a standard GEO mission, in contrast to a highly-inclined orbit configuration, this type of system is acceptable because the bias states capture the effects of scale factor and misalignment uncertainties, whose bias effects are nearly constant over an orbit. However, if the spacecraft implements a SNP yaw-steering profile as in a highly-inclined orbit configuration, the scale factor and misalignment uncertainties induce gyro biases that are proportional to the spacecraft angular rate and corrupt the attitude determination solution. In such a case, the standard six-state Kalman filter cannot be used without resulting in significantly degraded antenna-pointing performance.
One possible solution to this problem is to implement attitude sensors such as star trackers and perform periodic calibration slew maneuvers where the spacecraft is reoriented from one inertial attitude to another. In such a solution, the difference between the gyro-propagated attitude and the measured attitude can be used to estimate the gyro scale factor and misalignment errors. The estimates may then be used to compensate the gyro rates before they are processed using the standard six-state Kalman filter. The disadvantage of this approach is that the spacecraft cannot perform its mission while performing the necessary calibration maneuvers, which may be needed as frequently as once a month, thereby resulting in significant mission downtime.
Another solution is set forth in U.S. Pat. No. 6,298,288 (“the '288 patent”), issued on Oct. 2, 2001, in which a fifteen-state Kalman filter is used to estimate spacecraft attitude values, wherein the fifteen-state Kalman filter includes three states related to rate dependent gyro scale factor errors and six states related to rate dependent gyro misalignment errors. The fifteen-state Kalman filter disclosed in the '288 patent is applied to measured gyro rates ωm, in order to obtain corrected gyro rates ωc which are then used to propagate attitude estimates. Although the fifteen-state Kalman filter disclosed in the '288 patent attempts to account for gyro scale factor errors and gyro misalignment errors, such errors are not sufficiently filtered from the resultant corrected gyro rates because the use of measured gyro rates ωm in the Kalman filter propagates the gyro scale factor errors and the gyro misalignment errors through the Kalman filter to the “corrected” rates. Accordingly, the solution described in the '288 patent results in the output of noisy rates from the Kalman filter due to the effects of bias, noise, scale factor and alignment uncertainties that are inherent in measured gyro rates ωm, and will still be present to some degree even if “corrected” measured rates are used in the filter. Such noisy rates result in increased attitude estimation errors.
What is needed is an improved attitude determination system for yaw-steering spacecraft, such as those in a highly-inclined orbit configuration, that can automatically estimate and compensate for gyro scale factor and misalignment errors without interruption of the spacecraft mission, and without resulting in noisy rates and associated attitude estimate errors.